Nacelle flow assembly

ABSTRACT

A nacelle assembly for a turbine engine has a cowl for a turbine engine. The cowl has a first surface spaced from a second surface. The second surface defines defining a bypass flow passage. A flow volume is spaced between the first surface and the second surface. A plurality of holes are disposed on the cowl. Each of the plurality of holes are configured to alter local air pressure about one of the first surface and the second surface of the cowl. The plurality of holes are in communication with the flow volume.

RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.12/894,505 filed Sep. 30, 2010, which is a divisional of U.S.application Ser. No. 11/938,975 filed Nov. 13, 2007.

BACKGROUND OF THE INVENTION

This invention relates to a nacelle assembly for a gas turbine engine,particularly for an aircraft.

For a gas turbine engine, such as a turbo fan engine, air is pressurizedin a compressor and mixed with fuel in a combustor to generate hotcombustion gases. These gases flow downstream through the turbine stagesof the engine, which extract energy from the gases. In a two spool gasturbine engine, a high pressure turbine powers a high pressurecompressor, while a low pressure turbine powers the fan section disposedupstream of the compressor and a low pressure compressor.

Combustion gases are discharged from the turbo fan engine through a coreexhaust nozzle while fan air is discharged through an annular fanexhaust nozzle defined at least partially by a nacelle surrounding thecore engine. A majority of the propulsion thrust is provided by thepressurized fan air, which is discharged through the fan exhaust nozzle.The remaining thrust is provided by the combustion gases dischargedthrough the core exhaust nozzle.

It is known in the field of aircraft gas turbine engines that theperformance of the turbo fan engine varies during diverse flightconditions experienced by the aircraft. An inlet lip section located atthe foremost edge of the turbo fan nacelle is typically designed toenable operation of the turbo fan engine and prevent the separation ofairflow from the inlet lip section of the nacelle during these diverseflight conditions. For example, the inlet lip section requires a “thick”inlet lip section designed to support operation of the turbo fan duringspecific flight conditions, such as cross-wind conditions, take-off andthe like. Disadvantageously, the “thick” inlet lip section may reducethe efficiency of the turbo fan engine during cruise conditions of theaircraft, which exist for the large majority of the flight of theaircraft.

A need therefore exists to optimize the performance of a turbo fan gasturbine engine during diverse flight conditions so as to reduce thenacelle thickness and its associated drag.

SUMMARY OF THE INVENTION

A nacelle assembly for a turbine engine has a cowl for a turbine engine.The cowl has a first surface spaced from a second surface. The secondsurface defines defining a bypass flow passage. A flow volume is spacedbetween the first surface and the second surface. A plurality of holesare disposed on the cowl. Each of the plurality of holes are configuredto alter local air pressure about one or both of the first surface andthe second surface of the cowl. The plurality of holes are incommunication with the flow volume.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a side cross-sectional view of a gas turbine engine,including the present nacelle.

FIG. 2 illustrates a close up view of the nacelle assembly of FIG. 1,highlighting a plurality of holes and a flow volume.

FIG. 3 illustrates a view of the nacelle assembly of FIGS. 1 and 2,showing a plurality of holes on a first surface and a second surface ofthe assembly.

FIG. 4 illustrates a close up view of a pattern of the plurality ofholes of FIG. 3.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a gas turbine engine assembly 10, which includes “inserial flow communication” fan section 14, low pressure compressor 15,high pressure compressor 16, combustor 18, high pressure turbine 20, andlow pressure turbine 22. During operation, air is pressurized in thecompressors 15, 16 and mixed with fuel in the combustor 18 to generatehot combustion gases. The hot combustion gases flow through high and lowpressure turbines 20, 22, which extract energy from the hot combustiongases. The high pressure turbine 20 powers high pressure compressor 16through high speed shaft 19 while a low pressure turbine 22 powers fansection 14 and low pressure compressor 15 through low speed shaft 21.The invention is not limited to two spool axial gas turbine architecturedescribed and may be used with other architectures, such as a singlespool axial design and a three spool axial design.

As shown in FIG. 1, gas turbine engine assembly 10 is in the form of ahigh bypass ratio turbo fan engine mounted within nacelle assembly 38,in which most of the air pressurized by fan section 14 bypasses turbineengine 12 for the generation of propulsion thrust. Nacelle assembly 38includes fan cowl 42 and core cowl 30 within fan cowl 42. Fan cowl 42and core cowl 30 define outer flow surfaces of nacelle assembly 38. Airflow F₂ is received by nacelle assembly 38 and passes through fansection 14. Discharge air flow F1 is discharged from fan section 14through bypass flow passage 36, which is defined by fan cowl 42 and corecowl 30. Core exhaust gases C are discharged from turbine engine 12through a core exhaust nozzle 32, depicted schematically, definedbetween core cowl 30 and center plug 34 disposed coaxially therein alonga longitudinal centerline axis A of turbine engine assembly 10. Asshown, fan cowl 42 extends at least partially around axis A of turbineengine 12 and further extends along it.

Referring now to both FIG. 1 and FIG. 2, nacelle assembly 38 has firstsurface 46 and second surface 52 as shown. First surface 46 and secondsurface 52 define the outer aerodynamic surface of nacelle assembly 38.Second surface 52 further defines bypass flow passage 36. First surface46 is spaced from second surface 52 to create flow volume 54. Flowvolume 54 has interior surface 50, a surface spaced between firstsurface 46 and second surface 52.

Extending between flow volume 54 and each of first surface 46 and secondsurface 52 are holes 58. Each hole 58 is in air flow communication withflow volume 54. In addition, as shown in FIG. 3, plurality of holes 58extend radially around axis A as well as along axis A of nacelleassembly 38. With reference to FIG. 4, which shows a close-up of firstsurface 46, each hole 58 is spaced along axis A a distance S_(X) fromits neighboring hole 58. Moreover, each hole 58 is radially spaced adistance S_(R) from its neighboring radially spaced hole 58. (See alsoFIG. 3) S_(X) is between the range of 15 mm to 0.2 mm while S_(R) isbetween the range of 15 mm and 0.1 mm. Holes 58 can be any pattern orcombination of patterns.

Referring back to FIG. 2, flow volume 54 is divided into first chamber74 and second chamber 82 by wall 70. Both first chamber 74 and secondchamber 82 extend completely around axis A. First chamber 74 is in airflow communication with first device 90 and second device 94. Inaddition, second chamber 82 is in air flow communication with firstdevice 90 and second device 94. First device 90 is a blowing device thatprovides first air flow 78 in first direction Q while second device 94is a suction device that provides second air flow 86 in the direction ofsecond direction arrow T. First chamber 74 may accordingly have firstair flow 78 in first direction Q as provided by first device 90.Alternatively, first chamber 74 may have second air flow 86 in thesecond direction T as provided by second device 94. Likewise, secondchamber 84 may have first air flow 78 or, alternatively, second air flow86. Air flow through first chamber 74 and second chamber 82 areindependently and separately controlled by control unit 98, which is incommunication with sensor 102.

As will be explained, for a specific operable condition, as sensed bysensor 102, control unit 98 may choose to blow first air flow 78 throughfirst chamber 74 or suck second air flow 86 through this chamber.Separately, control unit 98 may blow first air flow 78 or suck secondair flow 86 through second chamber 82. Because of wall 70, air flow inone direction, say first direction Q, will not interfere with air flowin second direction, say second direction T, which is a directionopposite to the direction of arrow Q. Holes 58 permit either first airflow 78 or second air flow 86 to be communicated to first surface 46 andsecond surface 52. In this way, local air pressure may be increased ordecreased around first surface 46 and second surface 52, therebyaltering air flow around these surfaces of nacelle assembly 38.

For example, control unit 98 may direct first device 90 to blow firstair flow 78 in the direction of arrow Q out holes 58, out first area 106of first surface 46 while also directing second device 94, which is inair flow communication with second chamber 82, to suck air from secondarea 110 of second surface 52 and create second air flow 86 in thedirection of arrow T. In this way, one area, say first area 106, mayhave an increase of local air pressure around first surface 46 whileanother area, say second area 110, may have a decrease in local airpressure. In so doing, greater control over nacelle assembly 38 isestablished so that turbine engine assembly 10 can be optimized for theappropriate operable condition.

Accordingly, an operable condition of an aircraft is sensed by sensor102 and communicated to control unit 98. Control unit 98 may then selectbetween increasing local air pressure or decreasing local air pressureat each surface (first surface 46 and second surface 52) independentlyof the other surface. The operability conditions may be a static takeoffcondition, a crosswind condition, a climb condition, a cruise condition,a windmill condition or any other condition. Each of these conditionswill dictate a response by control unit 98 to make a selection or noselection at all. In this way, different areas of first surface 46 andsecond surface 52, such as first area 106 and second area 110, may havedifferent air flow through holes 58 to create different local airpressure conditions, thereby altering the aerodynamic configuration ofnacelle assembly 38 without changing its actual physical size.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that certain modifications would come within the scope of thisinvention. For that reason, the follow claims should be studied todetermine the true scope and content of this invention.

1. A method of altering air pressure on a surface of a nacelle assembly,the method comprising the steps of: (a) sensing an operable condition ofan aircraft; (b) selecting between increasing local air pressure anddecreasing local air pressure at the nacelle surface of the nacelleassembly through a plurality of holes on said nacelle surface based onthe sensed operable condition; and (c) decreasing air flow at a firstarea of the nacelle surface and increasing airflow at a second area ofthe nacelle surface, the first area different from the second area,wherein the first area is radially outwards of the second area.
 2. Themethod of claim 1 wherein the operable condition comprises at least oneof a static takeoff condition, a climb condition, a cruise condition,and a windmill condition.
 3. The method of claim 1 including the stepof: (d) increasing air flow at the nacelle surface to increase local airpressure at the nacelle surface.
 4. The method of claim 1 including thestep of: (d) decreasing air flow at the nacelle surface to decreaselocal air pressure at the nacelle surface.
 5. The method of claim 1including the steps of: passing air from a first interior volume of thenacelle to a first exterior area of the nacelle via first holes of theplurality of holes in the nacelle to increase local air pressure at thefirst exterior area of the nacelle; and passing air from a secondexterior area different from the first exterior area of the nacelle to asecond interior volume of the nacelle via second holes of the pluralityof holes in the nacelle to decrease local air pressure at the secondexterior area of the nacelle.
 6. The method of claim 1, wherein thesecond area of the nacelle surface defines a bypass flow passage of thenacelle assembly.
 7. The method of claim 1, wherein the plurality ofholes are uniformly radially spaced apart and uniformly axially spacedapart.
 8. The method of claim 1 including the step of: (d) passing airfrom a first interior volume of the nacelle to a first exterior area ofthe nacelle via first holes of the plurality of holes in the nacelle toincrease local air pressure at the first exterior area of the nacellewithout passing air from a second interior volume of the nacelle to asecond exterior area of the nacelle.
 9. A method of altering airpressure on a surface of a nacelle assembly, the method comprising thesteps of: (a) sensing an operable condition of an aircraft; and (b)selecting between increasing local air pressure and decreasing local airpressure at a nacelle surface of the nacelle assembly through aplurality of holes on said nacelle surface based on the sensed operablecondition; and (c) increasing air flow at a first area of the nacellesurface and decreasing airflow at a second area of the nacelle surface,the first area different from the second area, wherein the first area isradially inwards of the second area.
 10. The method of claim 9, whereinthe first interior volume is radially outwards of the second interiorvolume and separated by wall.